U.S. patent number 7,269,957 [Application Number 10/857,713] was granted by the patent office on 2007-09-18 for combustion liner having improved cooling and sealing.
Invention is credited to Vincent C. Martling, Zhenhua Xiao.
United States Patent |
7,269,957 |
Martling , et al. |
September 18, 2007 |
Combustion liner having improved cooling and sealing
Abstract
A gas turbine combustion liner is disclosed having an alternate
interface region between it and a transition duct where the cooling
effectiveness along the aft end of the combustion liner is
improved, resulting in extended component life, while utilizing a
simpler combustion liner geometry. The region of the combustion
liner proximate its second end comprises a plurality of spring
seals that seal against a transition duct while admitting a cooling
fluid to pass into a passage, formed between the combustion liner
and spring seals, that feeds a plurality of cooling holes located
in the combustion liner proximate the liner second end. Depending
on the cooling requirements, the cooling holes can be angled both
axially and circumferentially to maximize the cooling
effectiveness.
Inventors: |
Martling; Vincent C. (Jupiter,
FL), Xiao; Zhenhua (Palm Beach Gardens, FL) |
Family
ID: |
35423677 |
Appl.
No.: |
10/857,713 |
Filed: |
May 28, 2004 |
Prior Publication Data
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Document
Identifier |
Publication Date |
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US 20050262845 A1 |
Dec 1, 2005 |
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Current U.S.
Class: |
60/800;
60/752 |
Current CPC
Class: |
F23R
3/002 (20130101); F01D 9/023 (20130101); F23R
3/46 (20130101); Y02T 50/675 (20130101); Y02T
50/60 (20130101) |
Current International
Class: |
F23R
3/60 (20060101) |
Field of
Search: |
;60/722,752,754,796,800
;431/343 |
References Cited
[Referenced By]
U.S. Patent Documents
Foreign Patent Documents
Primary Examiner: Casaregola; L. J.
Claims
The invention claimed is:
1. A combustion liner for a gas turbine engine, said combustion
liner comprising: a centerline; a first end; a second end; a first
cylindrical portion proximate said first end; a second portion
fixed to said first portion and extending towards said second end;
a third cylindrical portion fixed to said second portion proximate
said second end, said third portion comprising: an inner liner
wall; an outer liner wall; a first spring seal adjacent to said
outer liner wall and having a first length and a plurality of first
axial slots, with each of said first axial slots having a first
width; a second spring seal adjacent to said first spring seal and
having a second length and a plurality of second axial slots, with
each of said second axial slots having a second width; and, a
plurality of cooling holes extending from said outer liner wall to
said inner liner wall.
2. The combustion liner of claim 1 wherein said second portion is
generally conical.
3. The combustion liner of claim 1 wherein said second portion is
generally cylindrical.
4. The combustion liner of claim 1 wherein said first length is
greater than said second length.
5. The combustion liner of claim 1 wherein said first width is
greater than said second width.
6. The combustion liner of claim 1 wherein said cooling holes have
a hole diameter between 0.015 and 0.125 inches.
7. The combustion liner of claim 1 wherein said cooling holes are
oriented generally towards said second end at a first angle .alpha.
relative to said outer wall.
8. The combustion liner of claim 7 wherein said first angle .alpha.
is between 10 and 75 degrees.
9. The combustion liner of claim 7 wherein said cooling holes are
further oriented in a generally circumferential direction at a
second angle .beta. relative to said centerline.
10. The combustion liner of claim 9 wherein said second angle
.beta. is up to 80 degrees.
11. The combustion liner of claim 1 wherein said first and second
spring seals are fixed to said combustion liner proximate said
second end of said combustion liner.
12. The combustion liner of claim 1 wherein a cooling fluid passes
through said first axial slots, into an annular plenum formed
between said outer liner wall of said combustion liner and said
first spring seal, and through said cooling holes.
13. A combustion liner for a gas turbine engine, said combustion
liner comprising: a centerline; a first end; a second end; a first
cylindrical portion proximate said first end; a second generally
conical portion fixed to said first portion and extending towards
said second end; a third cylindrical portion fixed to said second
portion proximate said second end, said third portion comprising:
an inner liner wall; an outer liner wall; a first spring seal
adjacent to said outer liner wall and having a first length and a
plurality of first axial slots, with each of said first axial slots
having a first width; a second spring seal adjacent to said first
spring seal and having a second length and a plurality of second
axial slots, with each of said second axial slots having a second
width; and, a plurality of cooling holes extending from said outer
liner wall to said inner liner wall, said cooling holes oriented
generally towards said second end at a first angle .alpha. relative
to said outer wall and in a generally circumferential direction at
a second angle .beta. relative to said centerline.
14. The combustion liner of claim 13 wherein said first length is
greater than said second length.
15. The combustion liner of claim 13 wherein said first width is
greater than said second width.
16. The combustion liner of claim 13 wherein said cooling holes
have a hole diameter between 0.015 and 0.125 inches.
17. The combustion liner of claim 13 wherein said first angle
.alpha. is between 10 and 75 degrees.
18. The combustion liner of claim 17 wherein said second angle
.beta. is up to 80 degrees.
19. The combustion liner of claim 13 wherein said first and second
spring seals are fixed to said combustion liner proximate said
second end of said combustion liner.
20. The combustion liner of claim 13 wherein a cooling fluid passes
through said first axial slots, into an annular plenum formed
between said outer liner wall of said combustion liner and said
first spring seal, and through said cooling holes.
Description
TECHNICAL FIELD
This invention relates in general to gas turbine engines and more
specifically to the cooling and sealing arrangement of the aft end
of a combustion liner.
BACKGROUND OF THE INVENTION
A gas turbine engine typically comprises at least one combustor,
which mixes air from a compressor with a fuel. This fuel and air
mixture combusts after being introduced to an ignition source. The
resulting hot combustion gases pass through the combustion system
and into a turbine, where the gases turn the turbine and associated
shaft. A gas turbine engine is most commonly used for either
propulsion for propelling a vehicle or harnessing the rotational
energy from the engine shaft to drive a generator for producing
electricity. Most land-based gas turbine engines employ a plurality
of combustors arranged in a can-annular layout around the engine.
Referring to FIG. 1, a representative land based gas turbine engine
10 of the prior art is shown in partial cross section. Gas turbine
engine 10 comprises an inlet region 11, an axial compressor 12, a
plurality of combustors 13, each in fluid communication with a
transition duct 14, which are in fluid communication with a turbine
15. The hot combustion gases drive the turbine, which turns shaft
17 before exiting through outlet 16. Shaft 17 is coupled to the
compressor, and for power generation, to an electrical generator
(not shown).
The operating temperatures of the combustors 13 are typically well
over 3000 degrees Fahrenheit, while the temperature limits of the
materials comprising combustors 13 are much lower. Therefore, in
order to maintain the structural integrity for continued exposure
to the hot combustion gases, combustors 13 are cooled, typically by
air from compressor 12. However, it is critical to only use the
minimal amount of cooling air necessary to lower the operating
metal temperatures of combustor 13 to within the acceptable range,
and not use more air than necessary nor allow any cooling air
leakage.
In order to maximize the efficiency of the gas turbine engine, it
is imperative to minimize any leakage of air from compressor 12
that is not intended for cooling combustors 13, such that all air
not intended for cooling, passes through combustors 13 and
undergoes combustion. Leakage areas are especially common between
mating components such as the interface region between combustor 13
and transition duct 14. Seals or tight tolerances between such
mating components are typically employed to overcome such leakages
that can reduce overall performance and efficiency. However, it is
also imperative to provide adequate cooling to an interface
region.
Examples of prior art seals and cooling designs for the interface
region between combustor 13 and transition duct 14 are disclosed in
U.S. Pat. Nos. 5,724,816 and 6,334,310. The '816 patent pertains to
a plurality of axial channels that are formed between an inner
member and an outer member and can be used to cool the aft end
section of a combustion liner where it interfaces with a transition
duct. An example of this configuration is shown in FIG. 2 where a
combustion liner is provided having a plurality of axial cooling
channels 18. The '310 patent pertains to an alternate manner to
cool this same region of a combustion liner and can be used in
conjunction with the prior art combustion liner shown in FIG. 2.
Specifically, a combustion liner includes an outer cooling sleeve
that contains a plurality of cooling holes 19 for supplying cooling
air to the region between the liner and the outer cooling sleeve.
The outer cooling sleeve includes a swaged end such that when the
outer cooling sleeve is welded to the combustion liner the stresses
imparted to the outer cooling sleeve by a transition duct are moved
away from the weld joint. Often times these combustion liners are
also accompanied by at least one spring seal for sealing against
the inner wall of a transition duct.
While each of these designs are directed towards providing adequate
cooling at the interface region of a combustion liner and
transition duct, improvements can be made such that cooling
effectiveness is improved, extending component life, while
simultaneously minimizing unnecessary cooling air leakage.
SUMMARY AND OBJECTS OF THE INVENTION
The present invention seeks to provide a combustion liner having an
alternate interface region between it and a transition duct where
the cooling effectiveness along the aft end of the combustion liner
is improved, resulting in extended component life, while
accomplishing this utilizing a simpler combustion liner geometry.
The combustion liner comprises a first end, a second end, and is
formed from three portions, with the third portion fixed proximate
the second end and comprising an inner liner wall, an outer liner
wall, a first spring seal adjacent to the outer liner wall, a
second spring seal adjacent the first spring seal, and a plurality
of cooling holes extending from the outer liner wall to the inner
liner wall. Each of the first and second spring seals contain a
plurality of axial slots with the slots preferably offset
circumferentially. The second spring seal is positioned over the
first spring seal to control the amount of cooling air passing
through the plurality of first axial slots. Cooling air then passes
over the outer liner wall and through the cooling holes, which are
preferably angled both axially and circumferentially to maximize
the length of the cooling hole, before providing cooling to the
inner liner wall.
It is an object of the present invention to provide a combustion
liner having an interface region with a transition duct that has
improved cooling effectiveness.
It is another object of the present invention to provide a
simplified geometry for the aft end region of a combustion
liner.
In accordance with these and other objects, which will become
apparent hereinafter, the instant invention will now be described
with particular reference to the accompanying drawings.
BRIEF DESCRIPTION OF DRAWINGS
FIG. 1 is a partial cross section of a gas turbine engine of the
prior art.
FIG. 2 is a perspective view of a portion of a prior art combustion
liner.
FIG. 3 is a perspective view of a combustion liner in accordance
with the preferred embodiment of the present invention.
FIG. 4 is a cross section of a combustion liner in accordance with
the preferred embodiment of the present invention.
FIG. 5 is a detailed cross section of a portion of a combustion
liner in accordance with the preferred embodiment of the present
invention.
FIG. 6 is a top view of a portion of a combustion liner in
accordance with the preferred embodiment of the present
invention.
DETAILED DESCRIPTION OF THE PREFERRED EMBODIMENT
The preferred embodiment of the present invention is shown in
detail in FIGS. 3-6. Referring now to FIG. 3, combustion liner 20
is shown in perspective view with emphasis on its aft region, which
interfaces with a transition duct similar to that of transition
duct 14 in FIG. 1. Combustion liner 20 comprises a first end 21, a
second end 22, and a centerline A-A. Located proximate first end 21
is a first portion 23 that is generally cylindrical in shape. Fixed
to first portion 23 and extending towards second end 22 is a second
portion 24. Depending on the volume of fuel and air and required
flow velocities through combustion liner 20, second portion 24 may
be either generally cylindrical or generally conical. FIG. 4
reflects a generally conical shape to second portion 24. A third
portion 25 is fixed to second portion 24 opposite first portion 23
and proximate second end 22. Third portion 25 is shown in greater
detail in FIG. 5 and comprises an inner liner wall 26 and outer
liner wall 27 in spaced relation to form a liner wall thickness
28.
Located adjacent outer liner wall 27 is a first spring seal 29 that
has a first length 30 and a plurality of first axial slots 31 with
each of first axial slots 31 having a first width 32. Located
adjacent first spring seal 29 is a second spring seal 33 that has a
second length 34 and a plurality of second axial slots 35 with each
of second axial slots 35 having a second width 36. First spring
seal 29 and second spring seal 33 are primarily used as a sealing
system when combustion liner 20 is installed in a transition duct.
It is preferred that first and second spring seals, 29 and 33, are
fixed to combustion liner 20 proximate second end 22 such that
potential damage to seals 29 and 33 during installation of
combustion liner 20 in a transition duct is minimized. Further
characteristics of the spring seal arrangement have first length 30
of first spring seal 29 greater than second length 34 of second
spring seal 33 and first width 32 being greater than second width
36 in order to accommodate the necessary circumferential movement
of the spring seals as they are compressed during installation in a
transition duct while not preventing the necessary flow of cooling
fluid through this region. However, second spring seal 33 is offset
circumferentially from first spring seal 29 in order to reduce the
leakage of cooling fluid through second spring seal 33.
The spring seals form a generally annular passage 38 around a
portion of outer liner wall 27. A cooling fluid, such as air,
enters passage 38 through first plurality of axial slots 31 in
first spring seal 29. Located in third portion 25 of combustion
liner 20 is a plurality of cooling holes 37 that extend from outer
liner wall 27 to inner liner wall 26. The cooling fluid from
passage 38 flows through cooling holes 37 for providing both
effective film cooling along inner liner wall 26 and convective
cooling throughout thickness 28. Depending on the amount of cooling
or area requiring cooling the quantity, location, and size of
cooling holes 37 can vary. Due to the drop in cooling fluid
pressure across cooling holes 37 the flow of cooling fluid 37 is
regulated by cooling hole size, not the dimensions of axial slots
31 and 35.
Referring now to FIGS. 5 and 6, in order to maximize the
effectiveness of the cooling fluid passing through cooling holes
37, the cooling holes are preferably oriented generally towards
second end 22 at a first angle .alpha. relative to outer liner wall
27. Depending on the amount of cooling required, first angle
.alpha. ranges between 10 and 75 degrees. To further enhance the
effectiveness of the cooling through cooling holes 37, the cooling
holes are further oriented in a generally circumferential direction
at a second angle .beta. relative to centerline A-A. Depending on
the amount of cooling required, second angle .beta. is up to 80
degrees while cooling holes 37 have a hole diameter between 0.015
and 0.125 inches.
The preferred embodiment of the present invention is advantageous
since it can be easily tailored to a variety of cooling
requirements while simultaneously eliminating the outer member of
the prior art configuration and reducing machining requirements.
The amount of cooling fluid admitted into passage 38 from first
axial slots 31 is controlled as desired by adjusting the pressure
drop across cooling holes 37 by altering the size of cooling holes
37. Furthermore, these adjustments to cooling hole sizes also
regulate the amount of cooling through wall thickness 28 and along
inner liner wall 26.
While the invention has been described in what is known as
presently the preferred embodiment, it is to be understood that the
invention is not to be limited to the disclosed embodiment but, on
the contrary, is intended to cover various modifications and
equivalent arrangements within the scope of the following
claims.
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